In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor and the fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the gas turbine, upon which the hot combustion gases impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1800-2100xc2x0 F.
Many approaches have been used to increase the operating temperature limit and service lives of the turbine blades and vanes to their current levels. For example, the composition and processing of the base materials themselves have been improved. Cooling techniques are used, as for example providing the component with internal cooling passages through which cooling air is flowed.
In another approach, the surfaces of the turbine blades and vanes are coated with aluminum-containing protective layers that protect the articles against the combustion gas, and in some cases insulate the articles from the temperature of the combustion gas. For example, overlay aluminide coatings may be used on the surfaces of the articles. A ceramic layer may also overlie the protective layer on the surface. The articles are thereby able to run cooler and be more resistant to environmental attack.
Although all of these approaches are effective to improve the performance of gas turbines, there is an opportunity for additional improvements to extend the operating temperatures and service lives of the gas turbine components. There is a need for improved protective coating systems that extend the capabilities of the turbine components even further. The present invention fulfills this need, and further provides related advantages.
The present invention provides a nickel-base superalloy article protected by a protective layer, and a method for its preparation. The article is particularly useful as a gas turbine blade or vane. The article has a prolonged life in the thermal cycling conditions found in aircraft engine operation, as compared with conventional articles.
A nickel-base superalloy article protected by a protective layer comprises an article substrate having a surface and comprising a nickel-base superalloy, and a protective layer on the surface of the article substrate. The article substrate is preferably in the shape of a gas turbine component, most preferably a gas turbine blade or a gas turbine vane. The protective layer comprises nickel, aluminum, and at least two modifying elements selected from the group consisting of zirconium, hafnium, yttrium, and silicon. It is preferred that one of the modifying elements is zirconium, and the additional modifying elements are selected from the group consisting of hafnium, yttrium, and silicon. Each of the modifying elements zirconium, hafnium, and silicon which is present is included in the protective layer in an amount of from about 0.1 to about 5 (more preferably from about 0.1 to about 3 and most preferably from about 0.1 to about 1) percent by weight of the protective layer. The modifying element yttrium, where present, is included in the protective layer in an amount of from about 0.1 to about 1 percent by weight of the protective layer. The protective layer is preferably an overlay coating such as a predominantly beta (xcex2) phase NiAl composition, but it may be an overlay coating such as MCrAlX or a diffusion aluminide. Optionally, a ceramic layer may overlie the protective layer.
The protective layer desirably comprises from about 20 to about 35 weight percent aluminum, nickel, the at least two modifying elements, and possibly other elements interdiffused from the substrate. The protective layer is from about 0.0005 inch to about 0.004 inch thick.
The article is preferably prepared with an overlay coating by providing a substrate having a surface and comprising a nickel-base superalloy, and depositing a protective layer onto the surface of the substrate by a physical vapor deposition process. The layer comprises aluminum, nickel, and at least two modifying elements selected from the group consisting of zirconium, hafnium, yttrium, and silicon. Each of the modifying elements zirconium, hafnium, and silicon which is present is included in the protective layer in an amount of from about 0.1 to about 5 (more preferably from about 0.1 to about 3 and most preferably from about 0.1 to about 1) percent by weight of the protective layer. The modifying element yttrium, where present, is included in the protective layer in an amount of from about 0.1 to about 1 percent by weight of the protective layer.
It has been determined that the addition of a combination of two or more of the modifying elements zirconium, hafnium, yttrium, and silicon provide an improvement to the furnace cycle testing life of the protected article that is greater than is predicted based on simple additive effects. Accordingly, the use of only one such modifying element is expressly excluded from the scope of the invention. Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.